Gas turbine engine with low stage count low pressure turbine

ABSTRACT

A gas turbine engine includes a core nacelle defined about an engine axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A gear train is defined along an engine axis. The gear train defines a gear reduction ratio of greater than or equal to about 2.3. A fan drive turbine along the engine axis which drives the gear train. The fan drive turbine includes three to six (3-6) stages. A fan is configured for rotation within the fan nacelle for operation at a fan pressure ratio less than about 1.45. A fan variable area nozzle is axially movable relative to said fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation. A high bypass gas turbine engine is also disclosed.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. patent application Ser.No. 14/190,429, filed Feb. 26, 2014, which was a continuation-in-part ofU.S. patent application Ser. No. 13/340,988, filed Dec. 30, 2011, nowU.S. Pat. No. 8,800,914, which was a continuation-in-part of U.S. patentapplication Ser. No. 12/131,876, filed Jun. 2, 2008, now U.S. Pat. No.8,128,021.

BACKGROUND

The present invention relates to a gas turbine engine and moreparticularly to an engine mounting configuration for the mounting of aturbofan gas turbine engine to an aircraft pylon.

A gas turbine engine may be mounted at various points on an aircraftsuch as a pylon integrated with an aircraft structure. An enginemounting configuration ensures the transmission of loads between theengine and the aircraft structure. The loads typically include theweight of the engine, thrust, aerodynamic side loads, and rotary torqueabout the engine axis. The engine mount configuration must also absorbthe deformations to which the engine is subjected during differentflight phases and the dimensional variations due to thermal expansionand retraction.

One conventional engine mounting configuration includes a pylon having aforward mount and an aft mount with relatively long thrust links whichextend forward from the aft mount to the engine intermediate casestructure. Although effective, one disadvantage of this conventionaltype mounting arrangement is the relatively large “punch loads” into theengine cases from the thrust links which react the thrust from theengine and couple the thrust to the pylon. These loads tend to distortthe intermediate case and the low pressure compressor (LPC) cases. Thedistortion may cause the clearances between the static cases androtating blade tips to increase which may negatively affect engineperformance and increase fuel burn.

SUMMARY

In a featured embodiment, a gas turbine engine includes a core nacelledefined about an engine axis. A fan nacelle is mounted at leastpartially around the core nacelle to define a fan bypass airflow pathfor a fan bypass airflow. A gear train is defined along an engine axis.The gear train defines a gear reduction ratio of greater than or equalto about 2.3. A fan drive turbine along the engine axis which drives thegear train. The fan drive turbine includes three to six (3-6) stages. Afan is configured for rotation within the fan nacelle for operation at afan pressure ratio less than about 1.45. A fan variable area nozzle isaxially movable relative to said fan nacelle to vary a fan nozzle exitarea and adjust a pressure ratio of the fan bypass airflow during engineoperation.

In another embodiment according to the previous embodiment, the fandrive turbine defines a pressure ratio that is greater than about five(5).

In another embodiment according to any of the previous embodiments, thefan drive turbine defines a pressure ratio that is greater than five(5).

In another embodiment according to any of the previous embodiments, thefan bypass airflow defines a bypass ratio greater than about ten (10).

In another embodiment according to any of the previous embodiments, thefan bypass airflow defines a bypass ratio greater than ten (10).

In another embodiment according to any of the previous embodiments, thegear train defines a gear reduction ratio of greater than or equal toabout 2.5.

In another embodiment according to any of the previous embodiments, thegear train defines a gear reduction ratio of greater than or equal to2.5.

In another embodiment according to any of the previous embodiments, afan variable area nozzle is axially movable relative to the fan nacelleto vary a fan nozzle exit area and to adjust a pressure ratio of the fanbypass airflow during engine operation. A controller is operable tocontrol the fan variable area nozzle to vary the fan nozzle exit areaand adjust the pressure ratio of the fan bypass airflow.

In another embodiment according to any of the previous embodiments, thecontroller is operable to reduce the fan nozzle exit area at a cruiseflight condition.

In another embodiment according to any of the previous embodiments, thecontroller is operable to control the fan nozzle exit area to reduce afan instability.

In another embodiment according to any of the previous embodiments, thefan variable area nozzle defines a trailing edge of said fan nacelle.

In another featured embodiment, a high bypass gas turbine engineincludes a core nacelle defined about an engine axis. A fan nacelle ismounted at least partially around the core nacelle to define a fanbypass airflow path for a fan bypass airflow. A gear train is definedalong an engine axis. The gear train defines a gear reduction ratio ofgreater than or equal to about 2.3. A fan drive turbine is rotatableabout the engine axis which drives the gear train. The fan drive turbineincludes three to six (3-6) stages. A fan section is configured foroperation at a fan pressure ratio less than about 1.45. The fan bypassairflow includes a bypass ratio greater than about (10).

In another embodiment according to the previous embodiment, the fandrive turbine is a three (3) stage turbine.

In another embodiment according to the previous embodiment, the fandrive turbine is a five (5) stage turbine.

In another embodiment according to the previous embodiment, the fandrive turbine is a six (6) stage turbine.

In another embodiment according to the previous embodiment, the fandrive turbine defines a pressure ratio that is greater than about five(5).

In another embodiment according to the previous embodiment, the geartrain defines a gear reduction ratio of greater than or equal to about2.5.

In another embodiment according to the previous embodiment, the fandrive turbine defines a pressure ratio that is greater than five (5).The fan defines a fan pressure ratio less than about 1.45. The geartrain defines a gear reduction ratio of greater than or equal to 2.5.

In another embodiment according to the previous embodiment, there arethree turbine rotors. The fan drive turbine being the most downstream ofsaid three turbine rotors.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently disclosed embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1A is a general schematic sectional view through a gas turbineengine along the engine longitudinal axis;

FIG. 1B is a general sectional view through a gas turbine engine alongthe engine longitudinal axis illustrating an engine static structurecase arrangement on the lower half thereof;

FIG. 1C is a side view of an mount system illustrating a rear mountattached through an engine thrust case to a mid-turbine frame between afirst and second bearing supported thereby;

FIG. 1D is a forward perspective view of an mount system illustrating arear mount attached through an engine thrust case to a mid-turbine framebetween a first and second bearing supported thereby;

FIG. 2A is a top view of an engine mount system;

FIG. 2B is a side view of an engine mount system within a nacellesystem;

FIG. 2C is a forward perspective view of an engine mount system within anacelle system;

FIG. 3 is a side view of an engine mount system within another frontmount;

FIG. 4A is an aft perspective view of an aft mount;

FIG. 4B is an aft view of an aft mount of FIG. 4A;

FIG. 4C is a front view of the aft mount of FIG. 4A;

FIG. 4D is a side view of the aft mount of FIG. 4A;

FIG. 4E is a top view of the aft mount of FIG. 4A;

FIG. 5A is a side view of the aft mount of FIG. 4A in a first slideposition; and

FIG. 5B is a side view of the aft mount of FIG. 4A in a second slideposition.

FIG. 6 shows another embodiment.

FIG. 7 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1A illustrates a general partial fragmentary schematic view of agas turbofan engine 10 suspended from an engine pylon 12 within anengine nacelle assembly N as is typical of an aircraft designed forsubsonic operation.

The turbofan engine 10 includes a core engine within a core nacelle Cthat houses a low spool 14 and high spool 24. The low spool 14 includesa low pressure compressor 16 and low pressure turbine 18. The low spool14 drives a fan section 20 connected to the low spool 14 either directlyor through a gear train 25.

The high spool 24 includes a high pressure compressor 26 and highpressure turbine 28. A combustor 30 is arranged between the highpressure compressor 26 and high pressure turbine 28. The low and highspools 14, 24 rotate about an engine axis of rotation A.

The engine 10 in one non-limiting embodiment is a high-bypass gearedarchitecture aircraft engine. In one disclosed, non-limiting embodiment,the engine 10 bypass ratio is greater than about six (6), with anexample embodiment being greater than about ten (10), the gear train 25is an epicyclic gear train such as a planetary gear system or other gearsystem with a gear reduction ratio of greater than about 2.3 and the lowpressure turbine 18 has a pressure ratio that is greater than about 5.In one disclosed embodiment, the engine 10 bypass ratio is greater thanten (10:1), the turbofan diameter is significantly larger than that ofthe low pressure compressor 16, and the low pressure turbine 18 has apressure ratio that is greater than 5:1. The gear train 25 may be anepicycle gear train such as a planetary gear system or other gear systemwith a gear reduction ratio of greater than about 2.5:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

Airflow enters the fan nacelle F which at least partially surrounds thecore nacelle C. The fan section 20 communicates airflow into the corenacelle C to the low pressure compressor 16. Core airflow compressed bythe low pressure compressor 16 and the high pressure compressor 26 ismixed with the fuel in the combustor 30 where is ignited, and burned.The resultant high pressure combustor products are expanded through thehigh pressure turbine 28 and low pressure turbine 18. The turbines 28,18 are rotationally coupled to the compressors 26, 16 respectively todrive the compressors 26, 16 in response to the expansion of thecombustor product. The low pressure turbine 18 also drives the fansection 20 through gear train 25. A core engine exhaust E exits the corenacelle C through a core nozzle 43 defined between the core nacelle Cand a tail cone 33.

With reference to FIG. 1B, the low pressure turbine 18 includes a lownumber of stages, which, in the illustrated non-limiting embodiment,includes three turbine stages, 18A, 18B, 18C. The gear train 22operationally effectuates the significantly reduced number of stageswithin the low pressure turbine 18. The three turbine stages, 18A, 18B,18C facilitate a lightweight and operationally efficient enginearchitecture. It should be appreciated that a low number of stagescontemplates, for example, three to six (3-6) stages. Low pressureturbine 18 pressure ratio is pressure measured prior to inlet of lowpressure turbine 18 as related to the pressure at the outlet of the lowpressure turbine 18 prior to exhaust nozzle.

Thrust is a function of density, velocity, and area. One or more ofthese parameters can be manipulated to vary the amount and direction ofthrust provided by the bypass flow B. The Variable Area Fan Nozzle(“VAFN”) 42 operates to effectively vary the area of the fan nozzle exitarea 44 to selectively adjust the pressure ratio of the bypass flow B inresponse to a controller C. Low pressure ratio turbofans are desirablefor their high propulsive efficiency. However, low pressure ratio fansmay be inherently susceptible to fan stability/flutter problems at lowpower and low flight speeds. The VAFN 42 allows the engine to change toa more favorable fan operating line at low power, avoiding theinstability region, and still provide the relatively smaller nozzle areanecessary to obtain a high-efficiency fan operating line at cruise.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 20 of the engine 10 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without the Fan Exit GuideVane (“FEGV”) system 36. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

As the fan blades within the fan section 20 are efficiently designed ata particular fixed stagger angle for an efficient cruise condition, theVAFN 42 is operated to effectively vary the fan nozzle exit area 44 toadjust fan bypass air flow such that the angle of attack or incidence onthe fan blades is maintained close to the design incidence for efficientengine operation at other flight conditions, such as landing and takeoffto thus provide optimized engine operation over a range of flightconditions with respect to performance and other operational parameterssuch as noise levels.

The engine static structure 44 generally has sub-structures including acase structure often referred to as the engine backbone. The enginestatic structure 44 generally includes a fan case 46, an intermediatecase (IMC) 48, a high pressure compressor case 50, a combustor case 52A,a high pressure turbine case 52B, a thrust case 52C, a low pressureturbine case 54, and a turbine exhaust case 56 (FIG. 1B). Alternatively,the combustor case 52A, the high pressure turbine case 52B and thethrust case 52C may be combined into a single case. It should beunderstood that this is an exemplary configuration and any number ofcases may be utilized.

The fan section 20 includes a fan rotor 32 with a plurality ofcircumferentially spaced radially outwardly extending fan blades 34. Thefan blades 34 are surrounded by the fan case 46. The core engine casestructure is secured to the fan case 46 at the IMC 48 which includes amultiple of circumferentially spaced radially extending struts 40 whichradially span the core engine case structure and the fan case 20.

The engine static structure 44 further supports a bearing system uponwhich the turbines 28, 18, compressors 26, 16 and fan rotor 32 rotate. A#1 fan dual bearing 60 which rotationally supports the fan rotor 32 isaxially located generally within the fan case 46. The #1 fan dualbearing 60 is preloaded to react fan thrust forward and aft (in case ofsurge). A #2 LPC bearing 62 which rotationally supports the low spool 14is axially located generally within the intermediate case (IMC) 48. The#2 LPC bearing 62 reacts thrust. A #3 fan dual bearing 64 whichrotationally supports the high spool 24 and also reacts thrust. The #3fan bearing 64 is also axially located generally within the IMC 48 justforward of the high pressure compressor case 50. A #4 bearing 66 whichrotationally supports a rear segment of the low spool 14 reacts onlyradial loads. The #4 bearing 66 is axially located generally within thethrust case 52C in an aft section thereof. A #5 bearing 68 rotationallysupports the rear segment of the low spool 14 and reacts only radialloads. The #5 bearing 68 is axially located generally within the thrustcase 52C just aft of the #4 bearing 66. It should be understood thatthis is an exemplary configuration and any number of bearings may beutilized.

The #4 bearing 66 and the #5 bearing 68 are supported within amid-turbine frame (MTF) 70 to straddle radially extending structuralstruts 72 which are preloaded in tension (FIGS. 1C-1D). The MTF 70provides aft structural support within the thrust case 52C for the #4bearing 66 and the #5 bearing 68 which rotatably support the spools 14,24.

A dual rotor engine such as that disclosed in the illustrated embodimenttypically includes a forward frame and a rear frame that support themain rotor bearings. The intermediate case (IMC) 48 also includes theradially extending struts 40 which are generally radially aligned withthe #2 LPC bearing 62 (FIG. 1B). It should be understood that variousengines with various case and frame structures will benefit from thepresent invention.

The turbofan gas turbine engine 10 is mounted to aircraft structure suchas an aircraft wing through a mount system 80 attachable by the pylon12. The mount system 80 includes a forward mount 82 and an aft mount 84(FIG. 2A). The forward mount 82 is secured to the IMC 48 and the aftmount 84 is secured to the MTF 70 at the thrust case 52C. The forwardmount 82 and the aft mount 84 are arranged in a plane containing theaxis A of the turbofan gas turbine 10. This eliminates the thrust linksfrom the intermediate case, which frees up valuable space beneath thecore nacelle and minimizes IMC 48 distortion.

Referring to FIGS. 2A-2C, the mount system 80 reacts the engine thrustat the aft end of the engine 10. The term “reacts” as utilized in thisdisclosure is defined as absorbing a load and dissipating the load toanother location of the gas turbine engine 10.

The forward mount 82 supports vertical loads and side loads. The forwardmount 82 in one non-limiting embodiment includes a shackle arrangementwhich mounts to the IMC 48 at two points 86A, 86B. The forward mount 82is generally a plate-like member which is oriented transverse to theplane which contains engine axis A. Fasteners are oriented through theforward mount 82 to engage the intermediate case (IMC) 48 generallyparallel to the engine axis A. In this illustrated non-limitingembodiment, the forward mount 82 is secured to the IMC 40. In anothernon-limiting embodiment, the forward mount 82 is secured to a portion ofthe core engine, such as the high-pressure compressor case 50 of the gasturbine engine 10 (see FIG. 3). One of ordinary skill in the art havingthe benefit of this disclosure would be able to select an appropriatemounting location for the forward mount 82.

Referring to FIG. 4A, the aft mount 84 generally includes a first A-arm88A, a second A-arm 88B, a rear mount platform 90, a whiffle treeassembly 92 and a drag link 94. The rear mount platform 90 is attacheddirectly to aircraft structure such as the pylon 12. The first A-arm 88Aand the second A-arm 88B mount between the thrust case 52C at casebosses 96 which interact with the MTF 70 (FIGS. 4B-4C), the rear mountplatform 90 and the whiffle tree assembly 92. It should be understoodthat the first A-arm 88A and the second A-arm 88B may alternativelymount to other areas of the engine 10 such as the high pressure turbinecase or other cases. It should also be understood that other framearrangements may alternatively be used with any engine case arrangement.

Referring to FIG. 4D, the first A-arm 88A and the second A-arm 88B arerigid generally triangular arrangements, each having a first link arm 89a, a second link arm 89 b and a third link arm 89 c. The first link arm89 a is between the case boss 96 and the rear mount platform 90. Thesecond link arm 89 b is between the case bosses 96 and the whiffle treeassembly 92. The third link arm 89 c is between the whiffle treeassembly 92 rear mount platform 90. The first A-arm 88A and the secondA-arm 88B primarily support the vertical weight load of the engine 10and transmit thrust loads from the engine to the rear mount platform 90.

The first A-arm 88A and the second A-arm 88B of the aft mount 84 forcethe resultant thrust vector at the engine casing to be reacted along theengine axis A which minimizes tip clearance losses due to engine loadingat the aft mount 84. This minimizes blade tip clearance requirements andthereby improves engine performance.

The whiffle tree assembly 92 includes a whiffle link 98 which supports acentral ball joint 100, a first sliding ball joint 102A and a secondsliding ball joint 102B (FIG. 4E). It should be understood that variousbushings, vibration isolators and such like may additionally be utilizedherewith. The central ball joint 100 is attached directly to aircraftstructure such as the pylon 12. The first sliding ball joint 102A isattached to the first A-arm 88A and the second sliding ball joint 102Bis mounted to the first A-arm 88A. The first and second sliding balljoint 102A, 102B permit sliding movement of the first and second A-arm88A, 88B (illustrated by arrow S in FIGS. 5A and 5B) to assure that onlya vertical load is reacted by the whiffle tree assembly 92. That is, thewhiffle tree assembly 92 allows all engine thrust loads to be equalizedtransmitted to the engine pylon 12 through the rear mount platform 90 bythe sliding movement and equalize the thrust load that results from thedual thrust link configuration. The whiffle link 98 operates as anequalizing link for vertical loads due to the first sliding ball joint102A and the second sliding ball joint 102B. As the whiffle link 98rotates about the central ball joint 100 thrust forces are equalized inthe axial direction. The whiffle tree assembly 92 experiences loadingonly due to vertical loads, and is thus less susceptible to failure thanconventional thrust-loaded designs.

The drag link 94 includes a ball joint 104A mounted to the thrust case52C and ball joint 104B mounted to the rear mount platform 90 (FIGS.4B-4C). The drag link 94 operates to react torque.

The aft mount 84 transmits engine loads directly to the thrust case 52Cand the MTF 70. Thrust, vertical, side, and torque loads are transmitteddirectly from the MTF 70 which reduces the number of structural membersas compared to current in-practice designs.

The mount system 80 is compact, and occupies space within the corenacelle volume as compared to turbine exhaust case-mountedconfigurations, which occupy space outside of the core nacelle which mayrequire additional or relatively larger aerodynamic fairings andincrease aerodynamic drag and fuel consumption. The mount system 80eliminates the heretofore required thrust links from the IMC, whichfrees up valuable space adjacent the IMC 48 and the high pressurecompressor case 50 within the core nacelle C.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

FIG. 6 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 7 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The disclosedembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

1. A gas turbine engine comprising: a core nacelle defined about anengine axis; a fan nacelle mounted at least partially around said corenacelle to define a fan bypass airflow path for a fan bypass airflow; agear train defined along an engine axis, said gear train defines a gearreduction ratio of greater than or equal to about 2.3; a fan driveturbine along said engine axis which drives said gear train, said fandrive turbine including three to six (3-6) stages; and a fan configuredfor rotation within the fan nacelle for operation at a fan pressureratio less than about 1.45, a fan variable area nozzle axially movablerelative to said fan nacelle to vary a fan nozzle exit area and adjust apressure ratio of the fan bypass airflow during engine operation.
 2. Theengine as recited in claim 1, wherein said fan drive turbine defines apressure ratio that is greater than about five (5).
 3. The engine asrecited in claim 1, wherein said fan drive turbine defines a pressureratio that is greater than five (5).
 4. The engine as recited in claim1, wherein said fan bypass airflow defines a bypass ratio greater thanabout ten (10).
 5. The engine as recited in claim 1, wherein said fanbypass airflow defines a bypass ratio greater than ten (10).
 6. Theengine as recited in claim 1, wherein said gear train defines a gearreduction ratio of greater than or equal to about 2.5.
 7. The engine asrecited in claim 1, wherein said gear train defines a gear reductionratio of greater than or equal to 2.5.
 8. The engine as recited in claim1, further comprising: a fan variable area nozzle axially movablerelative to said fan nacelle to vary a fan nozzle exit area and toadjust a pressure ratio of the fan bypass airflow during engineoperation; and a controller operable to control said fan variable areanozzle to vary the fan nozzle exit area and adjust the pressure ratio ofthe fan bypass airflow.
 9. The engine as recited in claim 8, whereinsaid controller is operable to reduce said fan nozzle exit area at acruise flight condition.
 10. The engine as recited in claim 8, whereinsaid controller is operable to control said fan nozzle exit area toreduce a fan instability.
 11. The engine as recited in claim 8, whereinsaid fan variable area nozzle defines a trailing edge of said fannacelle.
 12. A high bypass gas turbine engine comprising: a core nacelledefined about an engine axis; a fan nacelle mounted at least partiallyaround said core nacelle to define a fan bypass airflow path for a fanbypass airflow; a gear train defined along an engine axis, said geartrain defines a gear reduction ratio of greater than or equal to about2.3; and a fan drive turbine rotatable about said engine axis whichdrives said gear train, said fan drive turbine including three to six(3-6) stages; and a fan section configured for operation at a fanpressure ratio less than about 1.45, wherein said fan bypass airflowincludes a bypass ratio greater than about (10).
 13. The engine asrecited in claim 12, wherein said fan drive turbine is a three (3) stageturbine.
 14. The engine as recited in claim 12, wherein said fan driveturbine is a five (5) stage turbine.
 15. The engine as recited in claim12, wherein said fan drive turbine is a six (6) stage turbine.
 16. Theengine as recited in claim 12, wherein said fan drive turbine defines apressure ratio that is greater than about five (5).
 17. The engine asrecited in claim 12, wherein said gear train defines a gear reductionratio of greater than or equal to about 2.5.
 18. The engine as recitedin claim 12, wherein said fan drive turbine defines a pressure ratiothat is greater than five (5), said fan defines a fan pressure ratioless than about 1.45, and said gear train defines a gear reduction ratioof greater than or equal to 2.5.
 19. The engine as recited in claim 12,wherein there are three turbine rotors, with said fan drive turbinebeing the most downstream of said three turbine rotors.